(1) Field of the Invention
The present invention relates to the field of flight control systems for rotorcraft that modify the pitch of the blades of a main rotor serving to provide the rotorcraft at least with lift. The present invention relates more particularly to such flight control systems including a tactile signal warning device for warning a rotorcraft pilot that a demand has been made for excessive mechanical power to be supplied by a power plant of the rotorcraft for driving at least the main rotor.
(2) Description of Related Art
It should be recalled that rotorcraft are rotary wing aircraft in which at least lift is provided by at least one main rotor having a drive axis that is substantially vertical. In the specific context of a helicopter, the main rotor provides the rotorcraft not only with lift, but also with propulsion in any travel direction.
Rotorcraft are also commonly provided with an anti-torque device serving to guide them in yaw, such as at least one auxiliary rotor having a drive axis that is substantially horizontal. By way of example, such an auxiliary rotor is a tail rotor, or in the context of a high-speed helicopter having forward propulsion, it may be formed by a propulsive propeller.
The rotor(s) of the rotorcraft is/are driven in rotation by a power plant. The behavior in flight of the rotorcraft can be modified by a pilot of the rotorcraft causing the pitch of the blades making up the rotary wing of the rotor(s) to vary cyclically and/or collectively. More particularly, the pilot causes the blades to move about respective pitch variation axes extending in the main direction in which each blade extends.
The pilot of the rotorcraft may potentially be a human pilot generating manual flight commands and/or an autopilot generating automatic flight commands.
With reference more particularly to the main rotor, varying the blade pitch cyclically changes the attitude behavior of the rotorcraft, and more particularly selectively changes its behavior in pitching and/or in roll. Varying the pitch of the blades collectively leads to a change in the lift force supplied by the main rotor of the rotorcraft and/or enables the rotorcraft to be guided along the gravity axis.
In order to generate flight commands manually for varying the pitch of the blades of the main rotor, a human pilot makes use of main control linkages that are operated manually (as opposed to automatically) by the pilot moving flight control members. The main control linkages can pivot the blades about their respective pitch variation axes by means of a mechanism involving swashplates mounted both to move axially and to rock on a mast carrying the rotor. The swashplates comprise a top swashplate connected to the blades via pitch control rods and mounted to rotate on a bottom swashplate that is connected to the main control linkages.
Conventionally, in order to vary the collective pitch of the blades of the main rotor, the flight control member is typically arranged as a pitch lever for actuating a first main control linkage to cause the swashplates to move axially. In order to cause the blades of the main rotor to vary cyclically, the flight control member is typically arranged as a stick for actuating a second main control linkage for tilting the swashplates.
In general, the main control linkages make use of hydraulic servo-controls that drive the blades of the main rotor about their pitch variation axes on the basis of flight commands generated by the human pilot acting on the flight control members. The use of such servo-controls enables the pilot to pivot the blades of the main rotor about their respective pitch variation axes accurately and without requiring a large amount of force.
Nevertheless, it should be observed that in spite of the assistance provided by the servo-controls in driving pitch changes of the blades, the human pilot must still apply drive forces to the flight control member that are sufficient to overcome any friction of greater or lesser extent that persists among the various members making up the mechanical transmission constituting the main control linkage.
In other embodiments, the main control linkage may also include transmission that is both mechanical in part and also optical or electrical in part.
Under such circumstances, the movements of the flight control member may be detected by an optical or electrical sensor generating a signal representative of the current position of the control member. Such a signal is then transmitted to a computer suitable for generating control signals for hydraulic servo-controls that drive the blades of the main rotor about their respective pitch variation axes.
Such a flight control member may then advantageously be in the conventional form of pedals, a stick, or a lever of large operating travel, i.e. for example a free end of the control member can move through a distance that is considerable, being several tens of centimeters.
In another embodiment, it is also possible for the control member to be in the form of a miniature stick, also known as a “joystick”. Such a joystick then moves through small distances. More precisely, the free end of a joystick can move through an amplitude that is limited to a few centimeters.
In order to generate flight commands automatically, the rotorcraft has auxiliary control linkages making use of an autopilot, with one auxiliary control linkage being associated with each of the travel axes of the rotorcraft.
In a simplified arrangement of said auxiliary control linkages, the autopilot is used at least in certain modes of operation, commonly referred to as “basic modes”, to provide assistance in stabilizing travel of the rotorcraft respectively at least in pitching, in roll, and in yaw. In an improved arrangement of said auxiliary control linkages, the autopilot is also used in operating modes, commonly referred to as “higher modes”, and serving respectively to guide the rotorcraft along its various travel axes in pitching, in roll, in yaw, and along the gravity axis.
Depending on the equipment of the rotorcraft, the autopilot may be capable of performing at least said basic functional mode and possibly also the higher mode.
In order to cause the pitch of the blades of the main rotor to vary, the autopilot generates activation orders for actuators engaged on a given main control linkage. Such actuators commonly comprise an actuator known as the “trim actuator”, and an actuator commonly known as the “series actuator”.
For a given main control linkage, the trim actuator is typically connected to the flight control member in parallel with the main control linkage. The trim actuator is dedicated to moving the main control linkage with movements that may be considerable.
In a common embodiment, the trim actuator is commonly used to transmit a sensation of force to the pilot in response to the pilot causing the blades to pivot by acting on the flight control member. For this purpose, the trim actuator incorporates a force return system that generates a resisting force against a human moving the flight control member, which force depends on the current pivot position of the blades about their respective pitch variation axes.
The force return system commonly uses at least one trim actuator motor, or indeed a clutch mechanism and optionally resilient deformation means, e.g. organized as a spring. On being activated, the force return system causes the trim actuator to be anchored to the main control linkage, thus enabling the trim actuator to generate said resisting force against a human moving the flight control member.
The trim actuator is anchored to the main control linkage, in particular by means of a clutch mechanism serving selectively either to anchor the trim actuator to the main control linkage in a clutched position, or else to release the main control linkage from the engagement exerted by the trim actuator, when in a declutched position.
The human pilot commonly has available a function for inhibiting the operation of the force return system, known as a “trim release”, in order to release the trim actuator from being anchored to the main control linkage.
The trim actuator is anchored to the main control linkage in particular by activating the motor and/or, where appropriate, by using the clutch mechanism. When the trim actuator is in the anchored situation, rotation of the motor is caused to depend on identifying those flight commands that are issued by the human pilot via the flight control member so as to vary said resisting force, either directly, or where appropriate, by acting via the resilient deformation means.
Under such conditions, it is possible to distinguish between various types of trim actuator that are commonly classified depending on their structure and the ways in which they operate, in particular for the purpose of generating said resisting force.
In particular, the following types of trim actuator are known among others:
motorized friction trim actuators of passive type providing a resisting force that is constant and independent of the position of the flight control member; and
anchorable motorized trim actuators providing a variable resisting force against the main control linkage being moved by a human acting on the flight control member.
The resisting force opposed by an anchorable motorized trim actuator varies under the effect of activating the motor of the trim actuator. Activating said motor of the trim actuator makes it possible to vary the force gradient opposed by the trim actuator against human movement of the flight control member. The force gradient opposed by the trim actuator is determined in particular depending on the relative position between the anchoring position of the trim actuator on the main control linkage and the current position of the flight control member.
Furthermore, among anchorable motorized trim actuators, it is possible to distinguish between anchorable motorized trim actuators of passive type and anchorable motorized trim actuators of active type, depending on the ways used for generating said resisting force.
For an anchorable motorized trim actuator of passive type, said resisting force is produced via a prestressed spring that is placed under greater or lesser tension by activating the motor. The prestress of the spring provides piloting comfort by compensating for friction in the main control linkage.
With an anchorable motorized trim actuator of active type, said resisting force is produced directly by the motor of the trim actuator that opposes a resisting torque against the main control linkage being moved. The motor of the trim actuator also serves to provide said piloting comfort by compensating for the friction in the main control linkage by generating a resisting torque of given value.
The series actuator is typically placed in series in the main control linkage, being dedicated to moving the main control linkage with movements that are smaller and faster than those generated by the trim actuator.
Furthermore, the power plant of the rotorcraft has one or more fuel-burning engines, in particular turboshaft engines. The current operating rating of the power plant depends on a regulator unit applying various different regulation ratings that are identified relative to a nominal regulation rating, commonly referred to as the “all engines operable” (AEO) rating.
Regulating the operating rating of the power plant serves to avoid the engine(s) being degraded under the effect of excessive use being made of the capability of the power plant to supply the mechanical power required by the rotorcraft. Several limit criteria are taken into account by the regulation rating in order to avoid such excessive use.
Among such limit criteria, mention may be made in particular of the following:
a limit criterion concerning the temperature of the gas leaving the high pressure turbine of the turboshaft engine(s);
a limit criterion on the speed of the gas generator and/or of the free turbine driven by the gas leaving the high pressure turbine; and
a limit criterion on the torque as admitted to a main power transmission gearbox with which the rotor(s) is/are engaged in order to be driven.
In addition to the AEO rating, various specific operating ratings for the power plant are usually defined relative to stages of flight of the rotorcraft.
Among these specific regulation ratings within the AEO rating, mention may be made in particular of the following:
a maximum continuous power (MCP) rating defining the maximum power authorized for continuous use of the engine(s) in order to comply with the constraints imposed by said limit criteria;
a maximum takeoff power (TOP) rating defining the maximum power that is authorized for use of the engine(s) over a predefined duration that is defined to be sufficient to enable the rotorcraft to take off; and
a maximum transitional power (MTP) rating defining the maximum power that is authorized for use of the engine(s) during a transitional stage of changing the travel speed of the rotorcraft, in particular while the rotorcraft is accelerating.
Under such conditions, in the field of aviation, it is necessary to take account of the possibility of one of the engines of a power plant of a motor-driven aircraft failing. In the event of such a failure, the number of engines available for supplying the rotorcraft with the necessary mechanical power is reduced.
That is why one engine inoperative (OEI) ratings have been established for regulating the engines of a power plant in the event of one of them failing. In the event of one engine failing, at least one other available engine operating in OEI rating is capable of delivering the mechanical power needed for operating the rotorcraft during a predefined duration so as to enable the rotorcraft to continue flying temporarily in spite of one of its engines being unavailable.
Various OEI ratings are commonly established for various stages of flight of the rotorcraft, such as for example the following common OEI ratings:
a very short duration OEI (VSD-OEI) rating during which the still-operational engine(s) is/are capable individually of being used at a contingency rating for a short duration of the order of 30 seconds while the rotorcraft is taking off;
a short duration OEI (SD-OEI) rating during which the still-operational engine(s) is/are individually capable of being used at a contingency rating for a short duration of the order of 2 minutes to 3 minutes during an advanced stage of takeoff of the rotorcraft; and
a long-duration or continuous OEI (C-OEI) rating, during which the still-operational engine(s) is/are individually capable of being used at a maximum power for a duration that is long, and potentially unlimited.
The unit for regulating the operation of the power plant, such as for example a full authority digital engine control (FADEC) has a regulation command that is delivered by a control unit of the rotorcraft, such as for example an automatic flight control system (AFCS).
In this context, there arises the general problem of the human pilot of the rotorcraft monitoring potential excessive use of the capabilities of the power plant for supplying the mechanical power required by the rotorcraft.
It is known to monitor the mechanical power margins available from the power plant by using display means, on the basis of data provided by a first limitation instrument (FLI).
The values of margins that are available using various different limit criteria are collected by the FLI, and the most constricting margin compared with the various operating ratings authorized for the power plant is taken into account when displaying the available power margin and/or, in equivalent manner, the margin available for varying the pitch of the blades of the main rotor.
Nevertheless, it is useful to reduce the workload on a human pilot, whose attention is essentially occupied by monitoring the environment outside the rotorcraft. That is why warning devices have been developed that generate at least one tactile signal that can be perceived by the human pilot via a flight control member.
Such warning devices serve to warn the human pilot of a situation that is potentially dangerous with respect to making excessive use of the capabilities of the power plant for supplying the mechanical power required by the rotorcraft.
It is appropriate to obtain warning information for transmission to the pilot in tactile manner that is variable depending on the degree of urgency with which the pilot needs to take action to reduce the mechanical power needs that the power plant must supply.
It is also appropriate to optimize reasonably the extent to which parameters are taken into account that could reveal excessive use of the capabilities of the power plant for supplying the mechanical power required by the rotorcraft.
It is also appropriate to ensure that the techniques used by such warning devices do not give rise to any risk of generating sudden changes to the operating speed of the power plant.
By way of example, proposals have been made by the Applicant in a Document US 2010/123045, or as described in Documents U.S. Pat. No. 7,098,811 and EP 2 631 172 (Bell Helicopter Textron, Inc.), and U.S. Pat. No. 7,262,712 and US 2006/071817 (Safe Flight Instrument, Corp.), or indeed in US 2005/004721 (Einthoven, Pieter G.) to provide a tactile signal warning device on a pitch lever dedicated to manually controlling collective variation in the pitch of the blades of a main rotor of a rotorcraft in order to inform the pilot of an overload state of the engine(s) of the rotorcraft.
For that purpose, the respective values of various flight parameters of the rotorcraft are collected and are transmitted to a calculation unit that determines a current or an anticipated overload state of the engines of the power plant. In the event of there being such an overload state, the calculation unit causes the warning device to be activated in manners that vary depending on the degree with which the pilot needs to act in order to reduce the demand for power.
In order to determine the overload state of the engine(s), the flight parameters of the rotorcraft that are taken into consideration are selected from those that conventionally reveal the operating state of the engine(s), such as the flight parameters commonly used for characterizing the above-mentioned limit criteria in the context of regulating the engine(s).
According to Document U.S. Pat. No. 7,098,811, the warning device is made up of firstly of a shaker, and secondly of a spring cartridge engaged on the pitch lever and constrained by an electric motor that is used under the control of the calculation unit.
At a first emergency threshold, the calculation unit activates the electric motor by switching on the spring cartridge to generate a resisting force against the pilot moving the pitch lever. At a second emergency threshold, the calculation unit activates the shaker.
Said emergency thresholds are determined as a function of the values of various flight parameters of the rotorcraft as supplied by a health and usage monitoring system (HUMS).
Under the effect of being switched on, the spring cartridge is rated to an anchoring value that is regulated by the electric motor. The anchoring value is calculated iteratively depending on the current values of said flight parameters by varying flight conditions of the rotorcraft in accordance with an anticipated value for the mechanical power to be delivered by the power plant as calculated iteratively by a predicted algorithm depending on the values of the flight parameters transmitted by the HUMS and/or by a control unit.
According to Document U.S. Pat. No. 7,262,712, the warning device is constituted by a shaker that generates shaking of amplitude and frequency that vary depending on variation in the urgency with which the pilot needs to act on the behavior of the rotorcraft. The urgency with which the pilot needs to act is determined by the calculation unit depending on the values of various flight parameters supplied by the on-board instrumentation of the rotorcraft, such as values for the temperature of gas at the outlet from the high pressure turbine of the engines and values of the torque or the speed developed by the engine(s) depending on the flight stages of the rotorcraft.
In addition, and as described in Document EP 2 631 172, there is also known a method of using a device providing warning by means of tactile signals and fitted to a flight control member of a rotorcraft. Such a warning device thus comprises in particular:
a warning member comprising a trim actuator, the warning member generating a tactile sensation that can be perceived by the pilot via the control member;
a flight control unit serving to identify command data relating to the pilot requesting lift to be supplied by the main rotor, which data is deduced as a function of position data supplied to the control unit by a position sensor; and
a predictor unit that acts iteratively at a certain frequency to calculate a power margin relating to the power available from the engine by taking the difference between a predicted power and a power limit relating to the power to be supplied by the engine as a function of its current operating speed.
Nevertheless, the structure and the operating conditions for such warning devices acting by means of tactile signals are not fully satisfactory in the context of the general operating constraints of a rotorcraft, which are becoming more and more complex.
Such constraints relate in particular to how the various flight control linkages are organized for generating and transmitting flight commands seeking to move the blades of the rotor(s) of the rotorcraft about their respective pitch variation axes, with this applying in particular to the main rotor.